Powder propulsive for rockets or other self-propelled projectiles



rates 3,044,255 Patented July 17, 1962 3 044 255 POWDER PROPULSV FRRUCKETS 0R STEER SELF-PROPELLED PROJECTILES Michel Preconl, Paris,France, assignor to Societe Technique de Recherches industrielles etMecaniques S.T.R.I.M., Paris, France, a society of France Filed May 9,1955, Ser. No. 506,766 Claims priority, appiication France May 14, 19544 Claims. (Cl. Sti-35.6)

This invention relates to rockets and similar missiles wherein thecombustion of an explosive or combustible charge carried by the missileserves to generate a jet of exhaust gases which is discharged through anexhaust nozzle to provide a forward thrust operative toV propel themissile along its path.

It is a general object of the invention to provide an improvedpropulsive arrangement for a device of the type specified.. Anothergeneral object'is to provide an improved rocket construction.

A more specific object is the provision of an arrangement lwhereby theavailable space alloted to the vcombustible charge in a rocketpropulsive system can be more completely illed to capacity thanheretofore w-ithout interfering with a satisfactory combustion of Saidcharge; it is an object consequently toprovide a rocket capable of beingprovided with a more powerful propulsive charge than was heretoforepossible for a rocket of similar size. Another object is to improve theuniformity of combustion of the rocket charge and thereby impart la moreuniform thrust, with consequent increase in the accuracy of fire.

A further object is the provision of `an improved firing or detonatingarrangements for a rocket; a specific object is the provision of asingle firing device for both the front and rear combustible charges ofa two-charge rocket. Other objects include the provision of improvedcentering means for the rocket charges, and improvement of the flow ofcombustion gases through the rocket. An important object resides in theprovisionof a relative arrangement and design of a front combustiblecharge and a rear 'combustible charge in a rocket assembly, wherebycontrolled combustion of both charges if automatically obtained,improving the flight characteristics of the rocket on its path.

According to van important aspect of the invention, there is provided arocket including a front combustible charge having an axialduct therein,a rear combustible charge axially spaced from the front charge and meansdefining a continuous annular space around said rear charge for the freeow of combustion gases therearound l towards Va rocket exhaust, andmeans for tiring said charges, whereby owing to the provision of saidaxial duct in the front charge and said annular peripheral space aroundsaid rear charge, combustion will concurrently proceed radially outwardsfrom said axial duct in the front charge and radially inwards from saidannular space in the rear charge. "I'he above and further objects,features and advantages of the invention will appear as the descriptionproceeds. In the accompanying diagrammatic drawings which illustrate anexemplary embodiment of the invention, v

FIG. 1 is an axial cross section, withV parts broken away, of a rocketembodying improvements according to the invention, and

FIG.` 2 is a fragmentary sectional view, on an enlarged scale, of apreferred embodiment of one detail of the rocket -illustrated in FIG. l.Y

The exemplary rocket illustrated essentially comprises a headsection Thaving a terminal fuse T1; a propellent section P and an expandable tailstructure-E.

'Ihe propellent section P comprises two chambers. A front chamber C isdefined by a front cylindrical casing section V and a rear chamber C' isdefined by a rear cylindrical casing section V. The two casing sectionsV and V' are assembled together in any suitable way, such A as by thethreaded connection shown. Between the chambers C and C and in theregion of the assembly between the two casing sections, there is definedan intermediate chamber Co. As shown, the rear casing section V issomewhat restricted in diameter in the part where it is assembled withthe front casing section V; moreover the front chamber is slightlysmaller in diameter than the rear chamber. Designatingby D, D and D0 thediameters -of the front chamber, the rear chamber and the intermediatechamber respectively, the following relation therefore holds: D D D.

An improved explosive charge is provided according to the invention,comprising two blocks of explosive arranged for so-called compensatedcombustion. The front charge consists of a tubular block B formed withan axial duct 1 therethrough and coated with a shell of suitablecombustion inhibiting material around its periphery Z and its end faces3 and 4. The rear charge B1 is a solid cylinder having a diameter D"less than the rear chamber diameter D and formed with suitably proliledconvex end surfaces, such as the hemispherical surfaces 5 and 6 shown.

The outer diameter D" of the front charge B is substantially equal tothe inner diameter D of the front ,chamber C. Under these conditions itwill be observed that the combustion of the front charge B can beeffected only through the axial duct 1 and there ywill be no iiow ofdischarged combustion gases along the outer periphery of the charge.Thus the inhibitor shell surrounding the charge will not tend to beripped olf or partially separated from around the charge and safeoperation is promoted. The front charge B is separated from the warheadT by a transverse partition or bulwark 7 against which the forward endof the charge P is directly applied through the forward end 4 of thecombustion inhibitor shell. Owing to the resulting absence of anychamber in the front part of the charge in which pressure might buildup, the buildup of dangerously high pressures is prevented and safety isfurther enhanced.

Interposed Within the casing between the front and rear charges B landB' and extending substantially within the intermediate chamber C0 is aspacer structure comprising a spider 8 which may have substantially theshape of a cross in transverse section. Each arm of the spider is formedadjacent its forward end with radial enlargements or shoulders 8positioned by rearward abutment against the shoulders defined by thefront ends of the rear casing section V. interposed between the perpenidicular forward ends of the spider arms and the adjacent rear face ofthe front charge Bare a set of resilient shims vor spring washers 9. Therearwardly extending arms of spider 8 are arcuately contoured to conformwith the partspherical front end of the rear charge B so as to supportand position the latter. l

The front parts of the spider arms are formed with cut-outs radiallyinwards of the outer enlargements -8 thereof so'as todeiine a recesspositioned within the intermediate chamber CO and adapted to receivetherein a novel primer or detonator assembly 10 which in accord'- `ancewith the invention constitutes a common detonator for the entire'explosive charge of the rocket'. The detonator structure will be morefully described presently.

The rear lend yof rear charge B is Vsupported .and centeredfwithin arear, spider structurell also crossshaped in transverse section andhaving suitably arcuatecontoured front 4arm surfaces. A centering pin 16projects axially in a forward direction from the center of the rearspider into a complementary axial recess in the rear end of rear chargeB'. Similarly a centering pin 17 projects axially rearwards from thecenter of front spider S into a recess in the front end of rear chargeC. Both end recesses of the rear charge are coated withcombustion-inhibiting material.

The exhaust nozzle assembly for the rocket comprises a convergingsection 12 followed by a diverging outlet section 13 secured to section12 by a threaded connection. The converging section 12 is provided bythe rear casing V of Ithe rocket body and comprises a circular forwardsection -`12' followed `by a conical section 12 followed in turn by aninflected connecting or throat section 12". The arms of rear spider 11are rearwardly contoured to conform with the inner contour of theconverging nozzle section 12 so as to support and position said spidertherein.

Surrounding the exhaust nozzle structure is a cylindrical tail supportwhich constitutes a rear-ward extension of the main body of the rocketand supporting the retractible tail fins E pivoted thereto in anysuitable manner so shown.

The detonator charge mentioned above has a forwardly projecting part -10thereof extending into the axial duct in the front main charge B.imbedded in the midst of the detonator charge 10 is an electric sparkingor detonator head from which a cable conductor 14 extends out of thedetonator 10 and through the annular space between the outer surface ofrear charge B and the inner surface of rear casing section V. The cable14 is brought out of the rocket through the exhaust nozzle thereof andis supported in the throat section of the nozzle by a fran,gible sealingdisc or element 15. The sealing element 15 has la main transversesection and a cylindrical flange section which is clamped between themating surfaces of the converging nozzle section 12 and diverging nozzle-section 13. Preferably the frangible element 15 has an axial boss 15formed with an aperture through which the cable 14 is threaded. The boss15 increases the rigidi-ty of the wall r15 and results in a cleanerbreakage of the wall i5 on detonation of the charge, in that it causesthe breakage to occur substantially solely by shear action and withoutbending or radial tearing effects which otherwise would lead to anincomplete destruction of the element 15 and would allow residual partsthereof to remain in the nozzle throat and subsequently impede theoutrush of the exhaust gases.

A preferred construction of the shear Wall 15 is partly illustrated inthe enlarged cross section of FIG. 2. In this construction the element15 has a thin shear-able section of thickness e in kits radially outwardpart and a thick main section 15a provided by a rearwardly directed-boss of thickness e. In FIG. 2, 15jc designates the diameter on whichthe wall 15 is clamped between the adjacent parts 15b and 15C of thenozzle sections, and 15e designates the outer diameter of Itheenlargement or 'boss 15a. It will be noted that 15e is -only veryslightly smaller than 15f, lso that there is defined a thin shearsection which is comparatively very small both in radial extent and inaxial or thickness dimension, adjacent to the clamped outer section ofthe element 15. In one practical embodiment of the invention, thedifference between diameters 15e and 15]L was about ll mm., and theenlarged thickness e was about twice the shear thickness e. Suchconstruction results in a particularly clean breakage of the frangibleseal `15 and permits accurate calibration of the breaking pressure underwhich the seal will burst. Moreover, it will be noted that on breakageof the seal 15 all `around the circumference of the shear section(between diameters 15e and 151), the cylindrical surface of shoulder 15goutwardly defining the boss 15a will be immediately guided by 4theclosely adjacent wall of the rear nozzle section and quickly andcompletely expelled out of the nozzle.

It will be noted that the general construction of the rocket described,and the arrangement of the main explosive charges thereinis such as toeliminate the presi f ence of a forward combustion chamber within thefront charge B. This in turn eliminates the necessity of providing asupporting spider for supporting the front end of the front charge B, aswas generally required in prior rockets of comparable type. 4

In operation, energizing the electric firing device through cable 14fires the single detonator charge 110. The projecting appendage 10immediately sets lire to the front charge B in the inner peripheral areathereof surrounding the axial duct in the charge. The front charge Bthen proceeds to burn radially outwards around its axial duct in aperfectly smooth concentric manner so that it remains centered in thecasing throughout its combustion by its outer periphery and end facessurrounded by the inhibitor shell.

Simultaneously or practically so, the flame from the detonator charge isswept rearwards and sets fire to the rear charge B. In this instance thecombustion is initiated over the peripheral surface of the charge owingto lthe annular gap present between said surface and the wall of casingV, and proceeds radially inwards therefrom. Throughout the Icombustionthe rear charge remains centered by the action of the spiders 8 and 11and axial centering pins 16 and 17.

'It has been found that the :arrangement of front and rear chargesdisclosed herein achieves a remarkably uniform thrust throughout thecombustion period of both charges, and especially during the criticalinitial phase of the combustion. At the start of the combustion periodthe volume flow of gases emanating from the front charge is small owingto the reduced combustion area corresponding to the recessed axialregion of the front charge, having the relatively small diameter d. Theannular space 18 'around the rear charge is so dimensioned as to affordan appropriate flow section for these gases. Thus this annular space 18can be provided relatively small, thereby increasing the permissiblepower capacity of the missile for a given volume. As the combustionprogresses, working radially outwards from the axial duct in the frontcharge and radially inwards from the periphery of the rear charge, thevolume flow of gases from the front charge gradually increases and thegas flow from the rear charge concurrently decreases at a comparablerate. Since the flow area defined between the periphery of the rearcharge and the casing wall gradually increases as the combustionproceeds, the increasing rate of discharge of gases from the frontcharge can be adequately taken care of in spite of the small initialdimension of the gap 18, without creating dangerous or objectionablepressure surges liable to detract from the safety and/ or accuracy offire. Thus it will be seen that owing to the compensated arrangement oflthe charges according to the invention, uniform, safe and reliableoperation of the rocket is achieved while -at the same time attainingmaximum capacity and filling ratio, i.e. optimum utilization of theavailable space within the device.

The operation is further enhanced owing to the absence Aof axial fiow ofgases through the rear section of the rocket, except possibly during thefinal phase of the combustion.

As already mentioned, the diameter D of the rear charge is selected onlyslightly smaller than the diameter D of the front charge.

Assuming the effective lengths of the front and rear charges are equal,it is evident that in order for both charge to burn for the same periodof time, lthe condition is D=D"-d, wherev D and D" are the outerdameters of the rear Iand front charges respectively, and d is thediameter of the axial duct in the front charge. I have found itpreferable however to give the rear charge a slightly larger value thanthat satisfying the above equation, so that it will burn somewhat longerthan the front chargeand will remain centered by the centering pins 16and i7 to the last, and prevent flow of the gases along the axis of therear section of the rocket. While it 5 is true that after the frontcharge has been burned up, the residual axial plug of combustiblematerial still present in the rear rocket chamber is liable to vbeexpelled withou-t burning owing to the sharp drop in pressure which willoccur, still itis found that the loss' of the small amount of fuelrepresented by this resi-dual plug, is more than olset by the advantagethe arrangement brings with it.

Otherwise stated, the diameter of the above-mentioned residual plug issubstantially D-(.Dd), and its deliberate loss requires the provision ofa correspondingly larger-diameter rear charge; I have found however thatthe fuel material is utilized to much greater .advantage on theperiphery of the charge than axially. y

It will Ibe noted that the domed contour imparted t0 the ends of therear charge according to the invention greatly promotes the smoothnessof the peripheral How of gases throughV the device, minimizing friction,turbulence and the like.

The provision of the central chamber C0 in additio to the ladvantagesstemming from thesingle central type of detonation or firing provided bythe invention, also facilitates the assembly of the rocket using atwo-part casing with a threaded connection between the Vcasing parts inthe region of such central chamber.

What I claim is:

` 1. A rocket or similar self-propelled projectile comprising acylindrical casing having a front chamber, an intermediate chamber Ianda rear chamber, a cylindrical block of powderk forming 4the front chargelocated in said f front chamber having an axially extending channeltherein, means for limiting combustion of said front charge to thesurface of'said axial channel, a solid cylindrical block of powderlocated in said rear chamber and being yspaced from the internal surfaceof said rear chamber and igniting means for said front and rear chargeslocated in said intermediate chamber so that combustion of said frontcharge takes place through said axial channel exclusively and combustionof said rear charge takes place along its periphery. Y

y 2. A rocket according to claim 1 wherein `the interior diameter ofsaid front chamber is between the interior diameter of lsaid rearchamber and the interior diameter of said intermediate chamber. y

3. A rocket according to claim 1 wherein the diameter of the rear chargeis greater than the difference between the diameter Vof the front chargeand the diameter of the axial channel of said front charge.

4. A rocket or similar self-propelled projectile comprising acylindrical casing having a front chamber, an intermediate chamber and arear chamber, a cylindrical block gof powder forming the front chargelocated in said'front References Cited in the le of this patent i UNITEDSTATES PATENTS 487,628 A Lockhart Dec..6, 1892 1,151,258 Fischer Aug.24, 1915 1,994,490 Skinner Mar. 19, 1935 2,206,809 Denoix July 2, 19352,436,364 McDowell Feb. 17, 1948 2,455,015 Mace et al. Nov. 30, 19482,561,670 Miller July 24, 1951 2,605,607 Hickman Aug. 5, 1952 2,691,459Whitmore Oct. 12, 1954 2,793,492 Sage et al May 28, 1957 FOREIGN PATENTS831,496 France June 7, 1938 1,004,819 France Dec. 5, 1951 1,058,495France Nov. 4, 1953 France Apr. 7, 1954

